The present invention relates generally to gas turbine engines, and, more specifically, to turbine rotor blades therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the combustion gases in a high pressure turbine (HPT) for powering the compressor, and in a low pressure turbine (LPT) for powering an upstream fan in a aircraft turbofan engine application, or for powering an output shaft for marine and industrial applications.
The first stage turbine rotor blade first receives the hot combustion gases from the combustor and is therefore subject to the high temperature thereof. Accordingly, first stage turbine blades are formed of state-of-the-art superalloy metals which maintain strength in the hostile environment of the HPT for maximizing blade life in service.
Furthermore, each turbine blade is hollow and includes specifically configured cooling circuits therein which use a portion of air bled from the compressor for internally cooling the turbine blade during operation, as well as providing an external film of thermally insulating air from the spent air discharged from the turbine airfoil through various rows of film cooling holes.
Each blade includes an airfoil extending outwardly from a platform joined to a dovetail for individually mounting the blades in the perimeter of a supporting rotor disk. The airfoil has a generally concave pressure side and a generally convex suction side which extend radially in span from the root of the airfoil at the platform to its radially outer tip, and also extend axially in chord between opposite leading and trailing edges.
The cooling circuits found in the airfoil may have a myriad of configurations specifically tailored for cooling the different portions of the airfoil differently against the different heat loads from the combustion gases channeled over the opposite pressure and suction sides. The concave-convex configuration of the airfoil creates different velocity and pressure distributions over the surfaces thereof for maximizing efficiency of energy extraction from the combustion gases for rotating the supporting disk during operation.
Accordingly, the heat loads from the combustion gases vary from root to tip and from leading edge to trailing edge in complex three dimensional patterns, which in turn affect the local cooling requirements of the different portions of the airfoil.
Each airfoil has a maximum thickness immediately aft of the leading edge and tapers to a thin trailing edge. Each airfoil typically also includes a small extension of the pressure and suction sides at the tip of the airfoil which defines a squealer rib surrounding an open tip cavity extending from a tip floor which encloses the top of the internal cooling circuits.
The leading edge of each airfoil that first receives the hot combustion gases, the thin trailing edge of each airfoil, and the small squealer rib at the airfoil tip are differently configured, have different functions, and have problems specific to the configurations thereof for obtaining adequate cooling to ensure a long useful life of the turbine blade in service operation.
Modern turbine blades have undergone decades of development which has substantially increased their useful life to many years or thousands of hours of operation in a gas turbine engine without undesirable thermal distress which would limit their useful lives. However, the life of the turbine blade is nevertheless limited by local thermal distress in any region thereof notwithstanding the distress-free performance of the majority of the blade.
For example, the turbine blade tip is one region of the blade which is difficult to adequately cool over the desired long useful life of the blade. The squealer rib around the airfoil tip is provided as a local extension of the pressure and suction sides for minimizing the radial gap or clearance between the tip and the surrounding turbine shroud to minimize undesirable leakage of the combustion gases therethrough during operation. Since turbine blades are subject to occasional rubbing with the surrounding turbine shroud, the small squealer ribs reduce the adverse affects of tip rubbing while ensuring integrity of the tip floor which encloses the internal cooling circuit.
The squealer rib itself is solid material and relatively thin, and is bathed in the hot combustion gases that flow axially along the pressure and suction sides during operation, as well as radially over the pressure side, and through the tip gap with the surrounding turbine shroud as the gases leak over the tips during operation. The squealer ribs are therefore subject to heating from both their outboard sides and inboard sides within the tip cavity, as well as along their radially outermost edges. And, the tips are subject to the high centrifugal velocities of the airfoil tips during rotation, and the high velocity of the hot combustion gases which flow downstream thereover during operation.
Accordingly, the prior art is replete with various configurations for cooling turbine blade tips having different complexity, different performance, and different effectiveness in an operating engine over its intended long useful life.
Turbine blade tips typically include a plurality of tip holes extending perpendicularly through the tip floor for filling the tip cavity with spent cooling air from the internal cooling circuit. In this way, the spent cooling air opposes ingestion of the hot combustion gases in the tip cavity for improving tip cooling.
Furthermore, film cooling holes are typically found near the blade tip on the pressure side for creating film cooling over the pressure side squealer rib during operation. In both configurations, the spent cooling air provides local film cooling of the outboard and inboard surfaces of the squealer rib.
Since the spent cooling air is discharged through the tip holes under pressure, the air is discharged in discrete jets at high velocity perpendicular to the tip floor which limits the cooling effectiveness thereof. Accordingly, symmetrical diverging tip holes may be introduced through the tip floor for diffusing the discharged air to decrease the velocity thereof and correspondingly increase the pressure for enhancing cooling inside the tip cavity.
In yet another conventional tip cooling arrangement, cylindrical tip holes may be inclined through the tip floor for impingement cooling the inner or inboard surfaces of the squealer rib, particularly on the airfoil pressure side having the greatest heat load therein. However, since the tip holes have relatively small diameters they cannot be manufactured in the original casting of the blade itself, but must be formed by post-casting drilling. Drilling requires access to the tip floor without damaging the cast squealer ribs. To be effective, inclined tip holes must be positioned closely adjacent to the squealer rib, but the squealer rib would thereby interfere with the fabrication of the close tip holes.
Accordingly, the inclined tip holes must firstly be fabricated prior to formation of the squealer rib, and then the squealer rib must be fabricated which increases the difficulty and cost of manufacture and destroys the unitary nature between the squealer rib and main airfoil typically manufactured with superalloys that are directionally solidified or single crystal alloys.
Accordingly, it is desired to provide a turbine rotor blade having improved tip cooling.